In an aircraft such as, for example, an airplane, the fuselage is one of the large primary structural components in addition to the wings. With the exception of a few sections, it comprises a simple tube with largely constant diameter.
In aircraft, such as large-capacity aircraft, that are nowadays in use, only metallic composite materials are installed in the fuselage. In this case, the individual fuselage cells are assembled of preformed shells. The joining of the fuselage shells is realized with rivets. Reinforcing devices (frames) and devices for connecting the reinforcing devices (stringers) are arranged on the inner side of the fuselage by means of bonding and/or riveting for reinforcement purposes.
According to current trends, composite materials, such as composite fiber materials, will in the future be used as primary structure in a fuselage. These materials generally include two main components, namely an embedding matrix and reinforcing fibers. Composite fiber materials have a high specific strength, i.e., a high strength-to-weight ratio. Components including these materials therefore weigh less, but still have the same or even better properties than the corresponding metallic materials used so far.
Composite fiber material provides the advantage of having a high specific strength, i.e., a high strength and simultaneously a very low weight. Composite fiber materials significantly differ from metals in various aspects:
Fiber composites comprise different layers that are built up into a composite either in a lamination process (prepreg) or in an infusion process. In this case, more or less intense setting processes occur during the curing of the composites and result in a change in the thickness, as well as a shift of the layers relative to one another in curved components.
The properties of fiber composites result, among other things, from a largely undisturbed fiber structure. This means that fiber interruptions need to be minimized in order to optimally utilize the material properties.
Depending on the materials (textiles, prepregs) and processing methods used, very elaborate tools and time-consuming manufacturing processes are required.
The processing of composite fiber material components is very cost-intensive due to the required tools.
In contrast to a metallic skin panel, the layered structure may delaminate from outside under stress such that elaborate repairs are required.
Due to the light weight and the force paths that can be adapted in a customized fashion, however, composite fiber material may be the material of choice in future aircraft construction.
In the preparation of new cost-efficient manufacturing and installation concepts, the overall concept should be realized specific to the material. In addition, the process reliability of manufacturing processes for (large) components should be ensured. The repair of defective components not only involves high costs. The manufacturing processes, in particular, in the field of composite materials are relatively time-consuming Depending on the production figures, defective components or rejects need to be taken into consideration such that no production delays occur. Depending on the component and the structural size, this may require significant additional financial expenditures. The concept for the final assembly of all components should be simple, particularly for a large-scale manufacture, because a large portion of the overall manufacturing time is scheduled for the final assembly.
The first aircraft with a fuselage of composite fiber material, namely the Boeing 787, is composed of several wound fuselage cells that are manufactured in one piece in a winding process without fiber interruption. These fuselage cells are then assembled. In this type of manufacture, the tool used includes an elaborate internal core. This method is excellent with respect to the fiber structure because no interruptions are necessary.
However, there also are two disadvantages. On the one hand, an aircraft requires a very dimensionally accurate external geometry in order to not disturb the aerodynamics of the surface. An internal tool would be ideal for this purpose. However, this is not possible in winding processes such that other elaborate methods are used during the curing process in order to ensure the dimensional accuracy of the skin. In addition, the fibers inevitably become longer as the number of layers increases. During a setting process as it occurs during the curing of composite fiber materials, this may lead to a fiber surplus in the outer layers. Consequently, undulations may be created in the fiber layers.
When an internal tool is used, the skin is dimensionally accurate. However, this is associated with the problem that the inner fiber layers (i.e., the fiber layers toward the fuselage center) become shorter in a self-contained fuselage cell. These inner fiber layers cannot set when they are thermally acted upon and may lead to manufacturing concerns.
Both approaches, namely with an internal tool, as well as with an external tool, are optimal with respect to the fiber processing because no fiber interruptions exist, but can lead to the above-described problems. In addition, other objects, desirable features and characteristics will become apparent from the subsequent summary and detailed description, and the appended claims, taken in conjunction with the accompanying drawings and this background.